Abstract: A hydrogen fuelled aircraft includes a fuselage, a wing, at least one hydrogen fuelled gas turbine engine, at least one hydrogen fuel tank configured to store liquid or supercritical hydrogen and a hydrogen fuel line configured to delivery hydrogen fuel from the fuel tank to the engine. The gas turbine engine and the hydrogen fuel tank are each installed externally on the wing, and the hydrogen fuel line does not pass through the fuselage.

Abstract: A propulsive aircraft gas turbine engine comprises a turbine disposed in a gas turbine engine core flow and a recuperator heat exchanger disposed downstream of the turbine in gas turbine engine core flow, the recuperator heat exchanger being configured to transfer heat from the gas turbine engine core flow to gas turbine engine fuel. The recuperator heat exchanger is configured to accommodate a portion of the gas turbine engine core flow therethrough, the remainder being bypassed around the recuperator heat exchanger.

Abstract: A combustor assembly for gas turbine engine, combustor assembly including: combustor wall including interior surface and combustor wall opening, wherein interior surface partially defines combustion chamber, wherein combustor wall opening extends between combustion chamber and its exterior; sealing element disposed partially within combustor wall opening and includes air inlet, air outlet and air passageway, wherein air outlet exits into combustion chamber and delivers flow of air received from exterior of combustion chamber via air inlet and air passageway to combustion chamber; and fuel nozzle coupled to sealing element and configured to deliver fuel into combustion chamber, wherein combustor wall further includes first bearing surface and sealing element further includes second bearing surface, wherein first bearing surface is concave and forms part of first spherical surface such that first and second bearing surfaces are configured to move relative to each other about central point of first spherical sur

Abstract: An aircraft turbofan gas turbine engine includes a fan assembly, a compressor module, and a turbine module. An electric machine positioned downstream of the fan assembly is rotationally connected to the turbine module. The fan assembly is in fluid communication with the compressor module by an intermediate duct and includes a highest pressure fan stage having a plurality of fan blades defining a fan diameter. The compressor module includes a lowest pressure compressor stage having a row of rotor blades. An intermediate flow axis is defined joining a radially outer tip of a trailing edge of one of the fan blades of the highest pressure fan stage, and a radially outer tip of a leading edge of one of the rotor blades of a leading edge of a lowest-pressure compressor blade. An intermediate flow axis angle and the intermediate flow axis angle is from ?20 to ?30 degrees.

Abstract: A method of operating a gas turbine engine is disclosed, the gas turbine engine comprising a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor. The fuel management system comprises two fuel-oil heat exchangers through which oil and the fuel flow, the heat exchangers arranged to transfer heat to the fuel and comprising a primary fuel-oil heat exchanger and a secondary fuel-oil heat exchanger; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the two heat exchangers. The method comprises controlling the fuel management system so as to transfer between 200 and 600 kJ/m3 of heat to the fuel from the oil in the primary fuel-oil heat exchanger at cruise conditions.

Abstract: A compressor of an aircraft air pressurization system comprises: a rotor configured to be mechanically coupled to a spool of a gas turbine engine; an oil-lubricated bearing for supporting rotation of the rotor, wherein the oil-lubricated bearing is disposed within a bearing chamber; and a seal assembly for restricting oil from the bearing chamber from reaching an inlet of the blower compressor, wherein the seal assembly comprises: a seal disposed between the bearing chamber and an air buffer chamber; an air buffer inlet configured to receive a flow of pressurized air into the air buffer chamber; and an air buffer outlet configured to allow the flow of pressurized air to be exhausted from the air buffer chamber together with oil from the bearing chamber that has passed through the seal and become entrained within the flow of pressurized air.

Abstract: A method for determining one or more characteristics of a hole includes obtaining depth data, tool velocity data, process current data, and a plurality of gates. The method further includes determining a plurality of first, second, and third key points. The plurality of first, second, and third key points together form a plurality of key points. The method further includes obtaining a plurality of rules. Each rule includes one of: a relationship between a respective key point and a value associated with the respective key point; and a relationship between two or more respective key points. The method further includes determining the one or more characteristics of the hole based on the plurality of rules and the plurality of key points.

Abstract: A compressor of an aircraft air pressurization system comprises: a rotor configured to be mechanically coupled to a spool of a gas turbine engine; and a housing, wherein the rotor is supported for rotation within the housing about a rotor axis, wherein the rotor and housing define a primary air channel extending between an inlet of the blower compressor for receiving the inlet flow of air and an outlet from which the blower compressor is configured to output pressurized air to a delivery line. The housing further comprises a bleed opening in a wall of the housing between the inlet and the outlet, the bleed opening extending circumferentially about the rotor axis, wherein the bleed opening is for contaminants within the inlet flow of air, which have been driven outward due to the rotation of the rotor, to be bled out of the blower compressor.

Abstract: A method of analysing a non-contact excitation system for a spin rig includes, calculating a magnetic field per test piece element, integrating the magnetic field over the test piece element surface to get a force per test piece element, performing a Fourier analysis of the forces acting on the test piece element as it moves past the periodic array of exciter elements, calculating the damping added to the test piece element due to magnetic induction, calculating a vibration response of the test piece element as it moves past the stationary array of exciter elements, calculating a transient response of the test piece element where the stationary array of exciter elements is rapidly moved away from the test piece elements, and using the calculated transient response of the test piece element to calculate a required electromagnetic holding force for the array of exciter elements.

Abstract: A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

Abstract: A aircraft gas turbine engine and operation method, the engine including: a staged combustion system having pilot and main fuel injectors, and operates in a pilot-only range wherein fuel delivers to pilot fuel injectors, and a pilot-and-main operation range wherein fuel is delivered to at least the main fuel injectors. The engine further includes a fuel delivery regulator to pilot and main fuel injectors, which receives fuel from a first and second source containing fuels each with different characteristics. The staged combustion system switches between pilot-only and pilot-and-main range operation when in steady cruise mode, the mode defining a boundary between first and second engine cruise operation range. The fuel delivery regulator delivers fuel to pilot fuel injectors during at least part of the first engine cruise operation with different fuel characteristics from fuel delivered to one or both pilot and main fuel injectors the second engine cruise operation range.

Abstract: A method for shaping an aerofoil by: (a) defining an aerofoil having a nominal shape, the nominal shape defined by; a leading edge, a trailing edge, a root and a tip, a span extending from the root to the tip, a pressure surface and a suction surface extending from the leading edge to the trailing edge; a nominal camber line extending from the leading edge to the trailing edge; (b) defining an edge region on one of the pressure and/or suction surface which extends distance of at least 0.1% but no more than 10% of the camber line length from one of the leading edge or the trailing edge of the aerofoil; and (c) adapting the shape of the pressure and/or suction surface within the edge region such that the edge region of the aerofoil achieves an asymmetric profile with respect to the nominal camber line.

Abstract: A method of operating a gas turbine engine having an engine core having a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core; a fan shaft; a gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a heat exchange system having at least one fuel-oil heat exchanger arranged to transfer heat to the fuel; and a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located downstream of the at least one fuel-oil heat exchanger. The method includes operating the gas turbine engine using a fuel having a lubricity of between 0.71 mm and 0.90 mm wear scar diameter (WSD) at 25° C.

Abstract: There is provided a method of operating a gas turbine engine. The gas turbine engine comprises a staged combustor comprising an arrangement of fuel spray nozzles in which fuel flow is biased to a subset of the nozzles adjacent one or more ignitors during a re-light procedure. The method comprises providing fuel to the combustor having a calorific value of at least 43.5 MJ/kg. Also disclosed is a gas turbine engine.

Abstract: A method of operating a gas turbine engine, the engine including an engine core with a turbine, a compressor, a combustor arranged to combust a fuel, and a core shaft connecting the turbine to the compressor; a fan upstream of the engine core; a fan shaft; a main gearbox that receives an input from the core shaft and outputs drive to the fan via the fan shaft; a primary oil loop system to supply oil to lubricate the main gearbox; and a heat exchange system to transfer heat between the oil and the fuel, the oil having an average temperature of at least 180° C. on entry to the heat exchange system at cruise conditions. The method includes transferring heat from the oil to the fuel to lower the fuel viscosity to a value of less than or equal to 0.58 mm2/s on entry to the combustor at cruise conditions.

Abstract: Ternary resin blends which can be cured in order to form a cured resinous material for use as a matrix in composite materials. The ternary resin blends comprise a polyfunctional cyanate ester, an asymmetric bisphenol-derived cyanate ester, and a compound of formula (I) wherein R1, R2, R3, R4, R5 and R6 are as defined in the specification. A cured resinous material prepared by curing a ternary resin blend, a composite material comprising said cured resinous material, an article comprising said cured resinous material or said composite material, and methods for making the ternary resin blends, cured resinous materials, composite materials and articles are also disclosed.

Abstract: An electric machine includes a stator having a phase arrangement, and a rotor. The phase arrangement includes first and second legs connected in parallel at first and second primary junctions. The legs each include first and second coils connected in series through a first intermediate junction. The intermediate junctions are connected by a branch such that the phase arrangement is a bridge circuit. The phase arrangement permits an alignment current to flow between the primary junctions via the branch, the alignment current being conducted through an alignment current path passing through one coil of each leg. The phase arrangement includes a negative impedance converter to add a negative electrical impedance to the alignment current path by introducing additional electrical energy into the respective alignment current path. The alignment current causes a translational force to be applied to the rotor for maintaining alignment of the rotor with respect to the stator.

Abstract: A method of operating a gas turbine engine including: a combustor arranged to combust a fuel; and a fuel management system arranged to provide the fuel to the combustor, wherein the fuel management system includes two fuel-oil heat exchangers through which oil and fuel flow, which are arranged to transfer heat between the oil and fuel and include primary and secondary fuel-oil heat exchangers; a fuel pump arranged to deliver the fuel to the combustor, wherein the fuel pump is located between the heat exchangers; and a recirculation valve located downstream of the primary heat exchanger, the recirculation valve arranged to allow a controlled amount of fuel which has passed through the primary heat exchanger to be returned to the inlet. The method includes selecting one or more fuels such that the calorific value of the fuel provided to the gas turbine engine is at least 43.5 MJ/kg.

Abstract: An electric machine includes a stator having a phase arrangement, and a rotor. The phase arrangement includes legs connected in parallel at first and second primary junctions. Each leg includes a plurality of coils connected in series through a respective intermediate junction. Each leg is connected to at least one other leg by a branch at the respective intermediate junctions such that the phase arrangement is a bridge circuit. The phase arrangement conducts a motor current through the respective coils between the first and second primary junctions. The phase arrangement permits an alignment current to flow between the first and second primary junctions along an alignment current path which passes through at least one coil of two different legs via a respective branch. The motor currents cause a torque to be applied to the rotor, and the alignment current causes a translational force to be applied to the rotor.

Abstract: A temperature measurement system for a gas turbine engine, the gas turbine engine including, in axial flow sequence, a compressor section, a combustor section having plural fuel spray nozzles, and a turbine section. The temperature measurement system includes one or more optical thermometers, each optical thermometer configured to measure the temperature of a component washed by the working gas of the engine, the or each component being in the combustor section or the turbine section at a first position along the axis of the engine.